The present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
Gas turbine engines, such as those which power modern commercial and military aircrafts, include a compressor for pressurizing a supply of air, a combustor for burning a hydrocarbon fuel in the presence of the pressurized air, and a turbine for extracting energy from the resultant combustion gases. The combustor generally includes radially spaced apart inner and outer liners that define an annular combustion chamber therebetween. Arrays of circumferentially distributed combustion air holes penetrate multiple axial locations along each liner to radially admit the pressurized air into the combustion chamber. A plurality of circumferentially distributed fuel injector orifices axially project into a forward section of the combustion chamber to supply the fuel for mixing with the pressurized air.
Combustion of hydrocarbon fuel in the presence of pressurized air may produce nitrogen oxide (NOX) emissions that are subject to excessively stringent controls by regulatory authorities, and thus may be sought to be minimized. Lean-staged liquid-fueled aeroengine combustors can provide low NOx and particulate matter emissions, but are also prone to combustion instabilities. There are several mechanism that may cause combustion instabilities in radial-staged lean combustors including heat release concentrated in the front of the combustor, and weak flame holding at certain operating conditions where main stage air dilutes the pilot stage fuel-air ratio.